Airfoil with variable geometry expansion surface

ABSTRACT

A wing, or similar article of airfoil section, has a variable geometry surface for the active control of shock strength and transonic wave drag. In one embodiment, the wing has a region of distensible skin (4) aft of the line of maximum section, which extends along the span of the wing in those areas that experience drag. Pressure means (10, 20, 30) within the wing outwardly deflect the distensible region and produce a local bulge in the expansion surface. This bulge induces pre-shock compression and reduces the effect of the shock. The bulge is retracted by the natural elasticity of the skin material (which can be a conventional aluminum alloy) upon removal of the applied pressure. In another embodiment, the wing has a ramp portion (14) which is outwardly deflectable for the same purpose. The invention is applicable to supercritical and natural laminar flow wings.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to an article, such as an aircraft wing,having an aerofoil section with a variable geometry expansion surfacefor the control of shock strength.

The invention is particularly applicable to wings for transport aircraftwhere it is intended to combat transonic wave drag, to which such wingsare susceptible when the aircraft is flown in off-design conditions.However, the invention may also find use in the wings of other aircrafttypes as well as possibly for control surfaces rather than wings.

2. DISCUSSION OF PRIOR ART

It is well known that air-flows over aircraft wings at high subsonicfree stream Mach numbers exhibit shock waves due to localised excursionsinto supersonic flow. The supercritical wing forms of modern transportaircraft ensure that the effects of such shock waves are minimised atthe design cruise Mach number, altitude and coefficient of lift.However, it is recognised that these shock waves can increasesignificantly in strength with small departures from the designconditions, such as a change in the angle of incidence or an increase inMach number. These shock waves create drag, which is known as transonicwave drag, and can lead to the breakdown of ordered flow. Operationunder these conditions is obviously undesirable since it increasesaircraft fuel consumption. Moreover, the flow breakdown may causeproblems of aircraft control or reduction in aircraft structural lifethrough buffet loading.

In other, non-supercritical, wing sections designed for natural laminarflow, the aerofoil section necessary to maintain the laminar boundarylayer as far aft as possible (for the benefit in drag reduction thatthis conveys) produces transonic wave drag which unfortunately increasesrapidly in strength as either Mach number or lift coefficient rise abovethe design condition value. This places a severe limitation on theoperating band of Mach numbers and lift coefficients for laminar flowaircraft.

Various solutions have been proposed to alleviate the problem oftransonic wave drag associated with aircraft wings. One approach hasbeen to incorporate perforations or slots in the expansion surface ofthe wing at both the upstream and downstream sides of the shocklocation, with these being linked through a plenum chamber inside thewing. This arrangement makes use of the pressure differential thatexists in the airflow adjacent the wing. Air is bled from the higherpressure region downstream of the shock and reintroduced in jetsupstream of the shock. These jets create a ramp to oncoming flow andthereby cause compression waves which weaken the shock. One of thedrawbacks of this solution is that a boundary layer drag penalty isimposed across the entire range of flight conditions, due to thepresence of these holes or slots. This drag penalty may outweigh theadvantage gained by reduction of transonic wave drag in terms of overalloperating costs.

Another approach to reduction of transonic wave drag and associatedbuffeting is to configure the aerofoil section so as to produce the samepre-shock compression achieved by the method described above. Forexample, at page 673 of the Journal of Aircraft Volume 25, No 8, datedAugust 1988, Tai and co-workers disclose an aerofoil with a bulge on theexpansion surface at a position approximating to that of the shock wave.

However all aerofoil sections represent a compromise between variousconflicting requirements and it is unlikely that any specificconfiguration optimised with this specific end in view would have abeneficial or neutral effect across the whole range of flightconditions. It is more likely to be the case that, in securing areduction in the off-design point wave drag, some increase in viscous orwave drag is caused at another condition. Accordingly it is unlikelythat a wing of this nature would be of such advantage as to extend theflight envelope whilst reducing overall fuel consumption.

SUMMARY OF THE INVENTION

The present invention provides an aerofoil form with a variable geometryregion within its expansion surface which can be deployed to projectbeyond the normal outline of the aerofoil when in certain flightregimes, yet which can be withdrawn at other times to minimise boundarylayer drag.

The invention is a shock reducing device for an article of aerofoilsection, said device comprising means for varying the thickness of thearticle in comparison with its undeformed state, characterised in thatthe variation in thickness is confined to a region of the expansionsurface of the article, said region being centred aft of the line ofmaximum section of the article but forward of the trailing edge andextending along at least part of the span of the article, an outwarddeflection of said region serving to produce a shock reducing variationin the expansion surface of the article.

The variable geometry region is positioned ahead of the trailing edge(i.e. upstream of any trailing edge flaps or control surfaces) at achordwise position such that there is some outward displacement of theexpansion surface upstream of the position of the shock so as to causepre-compression of the flow ahead of the shock.

In some articles of aerofoil section, such as wings of laminar flowsection, the shock does not change chordal position significantly withchange in flight conditions so the variable geometry region can becentred on the expected position of the shock. For such a laminar flowaerofoil section, with the shock position at around 50% chord,satisfactory results have been obtained across a wide range of flightconditions using a variable geometry region extending from 45% to 65%chord. However, for a wing of the alternative supercritical section, theshock position can vary between 50% and 60% chord according to flightcondition. For such aerofoil sections, a variable geometry regionextending from 40% to 70% chord would cover the variation in shockposition, but for best results the pressure means would need to becapable of moving the centre of variation according to the anticipatedor detected position of the shock for a given flight condition, andactivated accordingly.

The variable geometry region can be incorporated into the article at anyspanwise location wherein the article is subject to shock whichincreases in strength with departure from design conditions. This can bein the inboard portion of the wing in certain modern transport aircraft(i.e. inboard of the trailing edge crank) but can also be the outboardportion of the wing in other aircraft or aircraft wing forms.

The degree of deflection required to combat transonic wave drag is quitesmall in relation to the thickness of the article. This is likely to be0.4% at most of the local chord. With this small degree of deflection asthe requirement the degree of elastic deformation demanded of thepertinent region of material might well be within the compass of currentmetal or reinforced plastics skinning materials when linked toappropriate pressure means.

In one form of the invention, the pressure means comprises apressurisable chamber within the article beneath a band of distensibleskin and means to cause pressurisation of the chamber to a degreesufficient to produce the required distension. This pressurisation maybe accomplished by pneumatic or fluidic means coupled to the aircraftcontrols or automatically activated by means of the aircraft air datasystem.

The invention may comprise mechanical pressure means as an alternativeto the pressurisable chamber discussed above. This can utilise one ofseveral forms of mechanical pushing device, such as cams or jacks.Several of these would be used in a co-ordinated manner to produce therequired distension of the skin. These mechanical Forms of pressuremeans have the advantage that they apply the distending pressure atparticular points on the skin. By actuating individual jacks or cams todifferent extents, the centre of distension can be adjusted toaccommodate variation in shock position such as that discussedpreviously in the context of supercritical wing sections.

It is believed that the precise form taken by the distensible skin whendeflected is not crucial to the success of the article in combatingtransonic wave drag. At present there seems to be no reason to departfrom a simple smooth curve profile. However, as the Mach number isincreased for a given lift coefficient, the normal shock wave may movedownstream. This depends on the particular aerofoil configuration underconsideration. Thus a distended skin having a crest at a fixed chordwiseposition may become less effective as the Mach number is increased. Infact, this approach can even make matters worse where buffet onset isconcerned. In the event that the requirement of adequate margin betweenbuffet onset Mach number and Mach number at cruise lift coefficientproves critical, a different approach may be necessary.

In an alternative form of the invention, the variable geometry regioncomprises a ramp which is actuable to a position where it projectsbeyond the undeformed profile of the article. Using such a device, thecrest of the deflected portion can be effectively moved downstream ofthe trailing edge. In its simplest manifestation, the ramp is a simplespoiler, which device is thought to be particularly effective inreducing buffet. However, calculations indicate that such devices conferminimal drag reduction due to the separation bubble which occursdownstream of the spoiler. An especially preferred form of ramp deviceincludes an integral fairing which closes off the region downstream ofthe ramp and minimises the degree of flow separation which occurs. Thispreferred form can also be designed to work in a spoiler mode, forexample during descent and landing when increased drag is desired, bydeploying itself as a single continuous plane.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will now by described by way of example with reference tothe drawings in which:

FIG. 1 shows a schematic representation of a wing in chordal section;

FIG. 2 is a part-sectional view of a first embodiment of the inventionshowing a system of cams effective to distend the wing skin;

FIG. 3 is an alternative embodiment which uses hydraulic jacks;

FIG. 4 gives a part-sectional view of the invention utilisingpressurisable chamber means;

FIG. 5 is a part-sectional view of another embodiment of the invention;

FIG. 6 is a part-sectional view of a preferred form of the inventionshown in FIG. 5, and

FIG. 7 is a data table and plot showing the effectiveness of aparticular bulge-form on a specified aerofoil configuration.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

The aerofoil section is depicted in all drawings using common referencenumerals for common parts. The article having aerofoil element comprisesan aircraft wing, generally designated 1, of supercritical form. Itsexpansion surface is designated 2 and the pressure surface designated 3.Within the skin of the expansion surface 2 there is a region 4 ofelastically distensible material which comprises one of the establishedaerospace alloys based upon aluminium, such as the aluminium-copperalloy designated AA2124. This region 4 is set into the expansion surfaceso that there is an absence of overlap ridges, in order to minimise skinfriction drag. The distensible region extends in a band along at leastpart of the span of the wing 1 over a chordal zone between 0.35 chordand 0.65 chord (measured from the leading edge), being centred at a linethrough 0.5 chord. This region 4 is configured and driven such that itproduces a bulge from the normal profile expansion surface 2 which, atits high point, projects outwardly by a distance of 0.25% wing chord.

The wing illustrated in FIG. 2 has a pressure means in the form of aseries of co-ordinated cams 10 driven by a common lead-screw 11. Thelead-screw 11 is rotated by a motor 12 to produce the required skindistension of region 4.

The wing illustrated in FIG. 3 has a pressure means in the form of aseries of co-ordinated jacks 20 supported by structural member 21. Thesemay be electrically or hydraulically driven.

An alternative version of the invention is depicted in FIG. 4. This formof wing has a pressurisable chamber 30 within the section of the wingand the material of region 4 is such that pressurisation of chamber 30causes the required degree of outward distension to produce the 0.25%chord bulge. In this embodiment, there is an internal feed pipe 31through which pressurising fluid is introduced to chamber 30 by means ofa pump (not shown).

In the embodiment depicted in FIG. 5, the region of variable geometrycomprises a ramp or spoiler device 14. As shown, this device is open atthe downstream end when deployed, allowing a bubble of essentiallytrapped air to form here. The benefit of reduced drag is not very greatowing to this Flow separation, but reduction in buffet excitation issignificant.

FIG. 6 shows an especially preferred form of the ramp device, in whichthe ramp 14 is complemented by a fairing 15 which closes off the regiondownstream of the ramp. This reduces flow separation to a negligiblelevel so that it no longer has a negative affect on the wave dragreduction achieved by the invention. The ramp 14 and fairing 15 arehinged at 16 and, as the ramp is moved outward into the flow, thetrailing edge of the fairing 15 is constrained to remain in contact withthe wing surface 2. In certain circumstances. For example during descentand landing, the ramp and fairing combination can be deployed in aspoiler mode in which they form a unitary continuous plane elementsimilar to the spoiler depicted in FIG. 5.

Actuation of such ramp devices is preferably performed by mechanicalmeans such as cams or jacks.

FIG. 7 records the effectiveness of a particular bulge geometry. Thedata shown in this Figure is modelling data secured by use of awell-proven two dimensional code. The plot of coefficient of pressureshows that the presence of the bulge decreases the magnitude of thepressure differential across the shock line and pushes this line in arearward direction. The table of data reveals that, for a commoncoefficient of lift (0.7) at a common Mach number (0.734), the bulgedconfiguration reduces the coefficient of drag from 0.0120 to 0.0105,i.e. a reduction of 13%. The coefficient of drag recorded hererepresents the overall drag experienced by the specified aerofoilsection. Thus, the noted decrease in overall drag indicates that thereduction in wave drag is not achieved at the expense of acounterproductive increase in viscous drag.

We claim:
 1. A shock reducing device for an article of airfoil section,said airfoil section having an expansion surface during transonicairflow conditions, said device comprising means for locally increasingthe steady state thickness of the article in comparison with saidarticle in an undeformed state, wherein said increase in thickness islocated in a region of the expansion surface of the article, said regionbeing centered aft of the line of maximum section of the article butforward of the trailing edge and extending along at least part of thespan of the article, wherein said increased steady state thickness insaid region comprising a means for producing a shock reducing variationin the expansion surface of the article.
 2. A device as claimed in claim1, wherein the article is an aircraft wing in which a region ofincreased steady static thickness occupies a chordal position betweenthe forward limit of 40% chord and the rearward limit of 70% chord.
 3. Adevice as claimed in claim 1, wherein the article is an aircraft wing ofa natural laminar flow section in which a region of increased steadystatic thickness extends from a forward chordal position at 45% chord toa rearward chordal position of 65% chord.
 4. A device as claimed inclaim 1, wherein the means for locally increasing the steady statethickness of the article is capable of moving a center of increasingthickness in a chordal direction.
 5. A device as claimed in claim 1,wherein a region of increased steady state thickness comprises adistensible skin portion and further includes a means for effectingdeflection of the distensible skin portion.
 6. A device as claimed inclaim 5, wherein the means for effecting deflection of the distensibleskin portion comprises an array of mechanical pushing devices.
 7. Adevice as claimed in claim 5, wherein the means for effecting deflectionof the distensible skin portion comprises a pressurizable chamber withinthe article together with a means for causing fluidic pressurization ofthis chamber.
 8. A device as claimed in claim 1, wherein the region ofvariable thickness comprises a ramp actuated by an array of mechanicalpushing devices.
 9. A device as claimed in claim 8, wherein the ramp hasan integral fairing which closes off the region downstream of the ramptrailing edge when the ramp is deployed.